Gas turbine engine buffer system

ABSTRACT

A gas turbine engine includes a buffer system that can communicate buffer supply air to a portion of the gas turbine engine. The buffer system can include a first circuit and a second circuit. The first circuit selects between a first bleed air supply having a first pressure and a second bleed air supply having a second pressure that is greater than the first pressure to render a first buffer supply air having an intermediate pressure. The second circuit selects between a third bleed air supply and a fourth bleed air supply to communicate a second buffer supply air.

BACKGROUND

This disclosure relates to a gas turbine engine, and more particularlyto a buffer system that can communicate a buffer supply air to one ormore portions of the gas turbine engine.

Gas turbine engines typically include at least a compressor section, acombustor section and a turbine section. During operation, air ispressurized in the compressor section and is mixed with fuel and burnedin the combustor section to generate hot combustion gases. The hotcombustion gases are communicated through the turbine section whichextracts energy from the hot combustion gases to power the compressorsection and other gas turbine engine modes.

Gas turbine engines typically include shafts that support a plurality ofairfoil supporting rotors of the compressor section and the turbinesection. Generally, these shafts are supported by bearing structuresthat define bearing compartments. The bearing compartments house one ormore bearings and contain lubricant that is used to lubricate thebearings. The lubricant is contained within the bearing compartment byone or more seals. A predetermined differential pressure must bemaintained across the seals so the lubricant cannot leak past the seals.

SUMMARY

A gas turbine engine includes a buffer system that can communicatebuffer supply air to a portion of the gas turbine engine. The buffersystem can include a first circuit and a second circuit. The firstcircuit selects between a first bleed air supply having a first pressureand a second bleed air supply having a second pressure that is greaterthan the first pressure to render a first buffer supply air having anintermediate pressure. The second circuit selects between a third bleedair supply and a fourth bleed air supply to communicate a second buffersupply air.

In a further embodiment of the foregoing gas turbine engine embodiment,each of the first circuit and the second circuit can include at leastone of an ejector and a valve.

In a further embodiment of either of the foregoing gas turbine engineembodiments, the ejector can be a variable area ejector that includes aplunger movably positioned within a nozzle to vary an orifice associatedwith the nozzle.

In a further embodiment of any of the foregoing gas turbine engineembodiments, one of the first circuit and the second circuit can includea conditioning device.

In a further embodiment of any of the foregoing gas turbine engineembodiments, the first buffer supply air is communicated to a componentsubject to a low pressure requirement of the gas turbine engine and thesecond buffer supply air is communicated to a component subject to ahigh pressure requirement of the gas turbine engine.

In a further embodiment of any of the foregoing gas turbine engineembodiments, the portion can include at least a bearing compartment ofthe gas turbine engine.

In a further embodiment of any of the foregoing gas turbine engineembodiments, the buffer system can include at least one controller incommunication with each of the first circuit and the second circuit.

In a further embodiment of any of the foregoing gas turbine engineembodiments, the controller can selectively control the communication ofthe first buffer supply air and the second buffer supply air.

In another exemplary embodiment, a gas turbine engine includes acompressor section, a combustor in fluid communication with thecompressor section, a turbine section in fluid communication with thecombustor, and a buffer system. The buffer system can include a firstcircuit that supplies a first buffer supply air and a second circuitthat supplies a second buffer supply air. The first circuit can includea first bleed air supply a second bleed air supply and at least one ofan ejector and a valve. The second circuit can include a third bleed airsupply, a fourth bleed air supply and at least one of an ejector and avalve.

In a further embodiment of the foregoing gas turbine engine embodiments,one of the first circuit and the second circuit can include aconditioning device.

In a further embodiment of any of the foregoing gas turbine engineembodiments, the ejector can be a variable area ejector.

In a further embodiment of any of the foregoing gas turbine engineembodiments, the gas turbine engine can include a high bypass gearedaircraft engine having a bypass ratio of greater than about six (6).

In a further embodiment of any of the foregoing gas turbine engineembodiments, the gas turbine engine includes a low fan pressure ratio ofless than about 1.45.

In a further embodiment of any of the foregoing gas turbine engineembodiments, the first circuit can include one of an ejector and a valveand the second circuit can include the other of the ejector and thevalve.

In yet another exemplary embodiment, a method of cooling a portion of agas turbine engine includes mixing a first bleed air supply with asecond bleed air supply to render a first bleed supply air of anintermediate pressure to the first bleed air supply and the second bleedair supply. The first buffer supply air can be communicated to acomponent subject to a first pressure requirement of the gas turbineengine. A second buffer supply air can be communicated to a componentsubject to a second pressure requirement of the gas turbine engine.

In a further embodiment of the foregoing method embodiment, thecomponent subject to the second pressure requirement can be separatefrom the component subject to the first pressure requirement.

In a further embodiment of either of the foregoing method embodiments, apower condition can be identified prior to the steps of communicating.

In a further embodiment of any of the foregoing method embodiments, thestep of mixing is performed in response to detecting a low powercondition during the step of identifying. The first bleed air supply canbe communicated as the first buffer supply air without performing thestep of mixing in response to detecting a high power condition.

In a further embodiment of any of the foregoing method embodiments, athird bleed air supply can be communicated as the second buffer supplyair in response to a high power condition of the gas turbine engine. Afourth bleed air supply can be communicated as the second buffer supplyair in response to a low power condition of the gas turbine engine.

In a further embodiment of any of the foregoing method embodiments, thesecond buffer supply air can be cooled prior to addressing the componentsubject to the second pressure requirement.

The various features and advantages of this disclosure will becomeapparent to those skilled in the art from the following detaileddescription. The drawings that accompany the detailed description can bebriefly described as follows.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 schematically illustrates a cross-sectional view of a gas turbineengine.

FIG. 2 illustrates a schematic cross-section of a portion of the gasturbine engine.

FIG. 3 illustrates an example buffer system that can be incorporatedinto a gas turbine engine.

FIG. 4 illustrates another example buffer system that can beincorporated into a gas turbine engine.

FIG. 5 illustrates yet another example buffer system that can beincorporated into a gas turbine engine.

FIG. 6 illustrates another exemplary buffer system.

FIG. 7 illustrates an example ejector of a buffer system, such as thebuffer system of FIG. 6.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates a gas turbine engine 20. The gasturbine engine 20 disclosed herein is a two spool turbofan engine thatgenerally incorporates a fan section 22, a compressor section 24, acombustor section 26 and a turbine section 28. Alternative engines mightinclude an augmenter section (not shown) among other systems orfeatures. The fan section 22 drives air along a bypass flow path B,while the compressor section 24 drives air along a core flow path C forcompression and communication into the combustor section 26. The hotcombustion gases generated in the combustor section 26 are expandedthrough the turbine section 28. Although depicted as a turbofan gasturbine engine in the disclosed non-limiting embodiment, it should beunderstood that the concepts described herein are not limited toturbofan engines and these teachings could extend to other types ofturbine engines, including but not limited to three spool enginearchitectures.

The gas turbine engine 20 generally includes a low speed spool 30 and ahigh speed spool 32 mounted for rotation about an engine centerlinelongitudinal axis A relative to an engine static structure 36 viaseveral bearing structures 38. It should be understood that variousbearing structures 38 at various locations may alternatively oradditionally be provided.

The low speed spool 30 generally includes an inner shaft 40 thatinterconnects a fan 42, a low pressure compressor 44 and a low pressureturbine 46. The inner shaft 40 can be connected to the fan 42 through ageared architecture 48 to drive the fan 42 at a lower speed than the lowspeed spool 30. The high speed spool 32 includes an outer shaft 50 thatinterconnects a high pressure compressor 52 and a high pressure turbine54. In this example, the inner shaft 40 and the outer shaft 50 aresupported at a plurality of points by bearing structures 38 positionedwithin the engine static structure 36. In one non-limiting embodiment,bearing structures 38 include at least a #1 bearing structure 38-1forward of the geared architecture 48 and a #2 bearing structure 38-2located aft of the geared architecture 48.

A combustor 56 is arranged between the high pressure compressor 52 andthe high pressure turbine 54. A mid-turbine frame 57 of the enginestatic structure 36 is arranged generally between the high pressureturbine 54 and the low pressure turbine 46. The mid-turbine frame 57 cansupport one or more bearing structures 38 in the turbine section 28. Theinner shaft 40 and the outer shaft 50 are concentric and rotate via thebearing structures 38 about the engine centerline longitudinal axis A,which is collinear with their longitudinal axes. The inner shaft 40 andthe outer shaft 50 can be either co-rotating or counter-rotating withrespect to one another.

The core airflow C is compressed by the low pressure compressor 44 andthe high pressure compressor 52, is mixed with fuel and burned in thecombustor 56, and is then expanded over the high pressure turbine 54 andthe low pressure turbine 46. The mid-turbine frame 57 includes airfoils59 which are in the core airflow path. The high pressure turbine 54 andthe low pressure turbine 46 rotationally drive the respective high speedspool 32 and the low speed spool 30 in response to the expansion.

In some non-limiting examples, the gas turbine engine 20 is ahigh-bypass geared aircraft engine. In a further example, the gasturbine engine 20 bypass ratio is greater than about six (6:1). Thegeared architecture 48 of the example gas turbine engine 20 includes anepicyclic gear train, such as a planetary gear system or other gearsystem. The example epicyclic gear train has a gear reduction ratio ofgreater than about 2.3. The geared architecture 48 enables operation ofthe low speed spool 30 at higher speeds which can increase theoperational efficiency of the low pressure compressor 44 and lowpressure turbine 46 and render increased pressure in a fewer number ofstages.

The low pressure turbine 46 pressure ratio is pressure measured prior toinlet of low pressure turbine 46 as related to the pressure at theoutlet of the low pressure turbine 46 prior to an exhaust nozzle of thegas turbine engine 20. In one non-limiting embodiment, the bypass ratioof the gas turbine engine 20 is greater than about ten (10:1), the fandiameter is significantly larger than that of the low pressurecompressor 44, and the low pressure turbine 46 has a pressure ratio thatis greater than about 5 (5:1). The geared architecture 48 of thisembodiment is an epicyclic gear train with a gear reduction ratio ofgreater than about 2.5:1. It should be understood, however, that theabove parameters are only exemplary of one embodiment of a gearedarchitecture engine and that the present disclosure is applicable toother gas turbine engines including direct drive turbofans.

In this embodiment of the example gas turbine engine 20, a significantamount of thrust is provided by a bypass flow B due to the high bypassratio. The fan section 22 of the gas turbine engine 20 is designed for aparticular flight condition—typically cruise at about 0.8 Mach and about35,000 feet. This flight condition, with the gas turbine engine 20 atits best fuel consumption, is also known as bucket cruise ThrustSpecific Fuel Consumption (TSFC). TSFC is an industry standard parameterof fuel consumption per unit of thrust.

Fan Pressure Ratio is the pressure ratio across a blade of the fansection 22 without the use of a Fan Exit Guide Vane system. The low FanPressure Ratio according to one non-limiting embodiment of the examplegas turbine engine 20 is less than 1.45.

Low Corrected Fan Tip Speed is the actual fan tip speed divided by anindustry standard temperature correction of “T”/518.7^(0.5). Trepresents the ambient temperature in degrees Rankine. The Low CorrectedFan Tip Speed according to one non-limiting embodiment of the examplegas turbine engine 20 is less than about 1150 fps (351 m/s).

FIG. 2 illustrates a portion 100 of a gas turbine engine, such as thegas turbine engine 20. The portion 100 can include one or more bearingstructures 38. Only one bearing structure 38 is depicted in FIG. 2 toschematically illustrate its features, but this is in no way intended tolimit this disclosure.

The bearing structure 38 supports a shaft 61, such as the inner shaft 40or the outer shaft 50, which supports a rotor assembly 63, such as arotor assembly of the compressor section 24 or the turbine section 28,through a hub 65. The rotor assembly 63 carries at least one airfoil 67for adding or extracting energy from the core airflow.

The bearing structure 38 defines a bearing compartment BC that housesone or more bearings 71. The bearing compartment BC contains a lubricantfor lubricating (and acting as a cooling medium to) the bearings 71. Oneor more seals 73 (two shown) contain the lubricant within the bearingcompartment BC. The seals 73 of the bearing compartment BC must bepressurized to prevent the lubricant from leaking out during certainground and flight conditions (both steady-state and transient). A buffersystem can be used to communicate buffer supply air to the bearingcompartment BC in order to provide adequate pressurization of the seals73 without exceeding material and/or lubricant temperature limitations.Example buffer systems that can be used for this and other purposes aredetailed below.

FIG. 3 illustrates an example buffer system 60 that can communicatebuffer supply air 62 to a portion of the gas turbine engine 20, such asto one or more bearing compartments BC. In this example, bearingcompartments BC-1, BC-2, BC-3, BC-4(a), BC-4(b) and BC-5 can be fed withbuffer supply air 62. The buffer supply air 62 pressurizes the bearingcompartments BC and can maintain the bearing compartments BC at anacceptable temperature. Although the example embodiment illustratescommunication of the buffer supply air 62 to multiple bearingcompartments BC-1 through BC-5 to provide adequate bearing compartmentseal pressurization to prevent lubricant leakage, buffer supply air 62could be communicated to only a single bearing compartment or could becommunicated for anti-icing, ventilation, cooling and other purposes.

The buffer system 60 includes a first bleed air supply 64 and a secondbleed air supply 66. In other words, the buffer system 60 is a dualsupply system. In the exemplary embodiment, the first bleed air supply64 is a low pressure bleed air supply and the second bleed air supply 66is a high pressure bleed air supply that includes a pressure that isgreater than the pressure of the first bleed air supply 64.

The first bleed air supply 64 can be sourced from the fan section 22,the low pressure compressor 44 or the high pressure compressor 52. Inthe illustrated non-limiting example, the first bleed air supply 64 issourced from an upstream stage of the high pressure compressor 52.However, the first bleed air supply 64 could be sourced from anylocation that is upstream from the second bleed air supply 66. Thesecond bleed air supply 66 can be sourced from the high pressurecompressor 52, such as from a middle or downstream stage of the highpressure compressor 52. The second bleed air supply 66 could also besourced from the low pressure compressor 44 or the fan section 22depending on from where the first bleed air supply 64 is sourced.

The buffer system 60 can also include a valve 68 that is incommunication with both the first bleed air supply 64 and the secondbleed air supply 66. Although shown schematically, the first bleed airsupply 64 and the second bleed air supply 66 can be in fluidcommunication with the valve 68 via buffer tubing, conduits, or otherpassageways. Check valves may also be used to prevent the second bleedair supply 66 from backflowing into the first bleed air supply 64.

The valve 68 can select between the first bleed air supply 64 and thesecond bleed air supply 66 to communicate the buffer supply air 62 to adesired portion(s) of the gas turbine engine 20. In other words, thebuffer supply air 62 that is communicated is either the first bleed airsupply 64 or the second bleed air supply 66 depending on which airsupply is ultimately selected by the valve 68, as is further discussedbelow.

The determination of whether to communicate the first bleed air supply64 or the second bleed air supply 66 as the buffer supply air 62 isbased on a power condition of the gas turbine engine 20. The term “powercondition” as used in this disclosure generally refers to an operabilitycondition of the gas turbine engine 20. Gas turbine engine powerconditions can include low power conditions and high power conditions.Example low power conditions include, but are not limited to, groundoperation, ground idle and descent idle. Example high power conditionsinclude, but are not limited to, takeoff, climb, and cruise conditions.It should be understood that other power conditions are alsocontemplated as within the scope of this disclosure.

In one exemplary embodiment, the valve 68 communicates the first bleedair supply 64 (which is a relatively lower pressure bleed air supply) asthe buffer supply air 62 in response to identifying a high powercondition of a gas turbine engine 20. The second bleed air supply 66(which is a relatively higher pressure bleed air supply) is selected bythe valve 68 and communicated as the buffer supply air 62 in response todetecting a low power condition of the gas turbine engine 20. Both thefirst bleed air supply 64 and the second bleed air supply 66 areintended to maintain the same minimum pressure delta across the bearingcompartment seals. Low power conditions require a higher stage pressuresource to contain the lubricant within the bearing compartment, whilehigh power conditions require a lower stage pressure source. The buffersystem 60 can use the lowest possible compressor stage to meet pressurerequirements in order to minimize supply temperature and any performanceimpact to the gas turbine engine 20.

The valve 68 can be a passive valve. A passive valve operates like apressure regulator that can switch between two or more sources withoutbeing commanded to do so by a controller, such as an engine control(EEC). The valve 68 of this example uses only a single input which isdirectly measured to switch between the first bleed air supply 64 andthe second bleed air supply 66.

The valve 68 could also be a controller based valve. For example, thebuffer system 60 can include a controller 70 in communication with thevalve 68 for selecting between the first bleed air supply 64 and thesecond bleed air supply 66. The controller 70 is programmed with thenecessary logic for selecting between the first bleed air supply 64 andthe second bleed air supply 66 in response to detecting a pre-definedpower condition of the gas turbine engine 20. The controller 70 couldalso be programmed with multiple inputs.

In one example, a sensor 99 detects a power condition of the gas turbineengine 20 and communicates a signal to the controller 70 to commandmodulation of the valve 68 between the first bleed air supply 64 and thesecond bleed air supply 66. The valve 68 could also be modulated to anintermediate level to inter-mix the first bleed air supply 64 and thesecond bleed air supply 66. Of course, this view is highly schematic. Itshould be understood that the sensor 99 and the controller 70 can beprogrammed to detect any power condition. Also, the sensor 99 can bereplaced by any control associated with the gas turbine engine 20 or anassociated aircraft. Also, although shown as a separate feature, thecontroller functionality could be incorporated into the valve 68.

FIG. 4 illustrates another example buffer system 160 that cancommunicate buffer supply air 162 to provide adequate bearingcompartment seal pressurization at an acceptable temperature. The buffersupply air 162 can also be used for additional purposes such asanti-icing and ventilation or for other cooling requirements of the gasturbine engine 20.

The buffer system 160 includes a first bleed air supply 164, a secondbleed air supply 166 and an ejector 172. If necessary, the first bleedair supply 164 can be augmented by the ejector 172 to prepare the buffersupply air 162 for communication to a portion of the gas turbine engine20, such as a bearing compartment BC (schematically shown by FIG. 4). Inother words, the ejector 172 can add pressure (using a relatively smallamount of the second bleed air supply 166) to the first bleed air supply164 to prepare the buffer supply air 162 for communication to anappropriate location of a gas turbine engine 20. In one exemplaryembodiment, the ejector 172 can mix the first bleed air supply 164 of afirst pressure with the second bleed air supply 166 of a second higherpressure to render the buffer supply air 162 of an intermediate pressureto the first bleed air supply 164 and the second bleed air supply 166.

The second bleed air supply 166, which is a higher pressure air than thefirst bleed air supply 164, can be communicated to the ejector 172 topower the ejector 172. The first bleed air supply 164 can be sourcedfrom the fan section 22, the low pressure compressor 44 or the highpressure compressor 52. The second bleed air supply 166 can be sourcedfrom a middle or downstream stage of the high pressure compressor 52, orcan include diffuser air. The second bleed air supply 166 could also besourced from the low pressure compressor 44 or the fan section 22depending on from where the first bleed air supply 164 is sourced.

Augmentation of the first bleed air supply 164 prepares the buffersupply air 162 at an adequate pressure and temperature to pressurize thebearing compartment(s) BC. The determination of whether or not toaugment the first bleed air supply 164 with the ejector 172 is based ona power condition of the gas turbine engine 20. Gas turbine engine powerconditions can include low power conditions and high power conditions.Example low power conditions include, but are not limited to, groundoperation, ground idle and descent idle. Example high power conditionsinclude, but are not limited to, takeoff, climb, and cruise conditions.It should be understood that other power conditions are alsocontemplated as within the scope of this disclosure.

In one example, the first bleed air supply 164 is augmented by theejector 172 in response to detecting a low power condition of the gasturbine engine 20 in order to communicate a buffer supply air 162 havingadequate pressurization. The amount of augmentation performed on thefirst bleed air supply 164 can vary depending upon the type of powercondition that is detected and the pressure requirements of the bearingcompartment(s) BC. For example, in one embodiment, the first bleed airsupply 164 is not augmented by the ejector 172 in response to detectionof a high power condition of the gas turbine engine 20. In other words,the first bleed air supply 164 can be communicated as the buffer supplyair 162 without any augmentation in response to some power conditions.

The buffer system 160 can include a controller 170 in communication withthe ejector 172 for determining whether or not to augment the firstbleed air supply 164. The controller 170 is programmed with thenecessary logic for making this determination in response to detecting apre-defined power condition of the gas turbine engine 20. In oneexample, a sensor 199 detects a power condition of the gas turbineengine 20 and communicates a signal to the controller 170 to command theejector 172 to augment the first bleed air supply 64. Of course, thisview is highly schematic. It should be understood that the sensor 199and the controller 170 can be programmed to detect any power condition.Also, the sensor 199 can be replaced by any control associated with thegas turbine engine 20 or an associated aircraft. Also, although shown asa separate feature, the controller 170 functionality could beincorporated into the ejector 172.

FIG. 5 illustrates yet another example buffer system 260. In thisexample, the buffer system 260 is a two-circuit, multi-source buffersystem that includes at least a first circuit 274 and a second circuit276. Additional circuits could also be incorporated. Low pressurerequirements of the gas turbine engine 20 can be fed with a first buffersupply air 262A from the first circuit 274, while high pressurerequirements of the gas turbine engine 20 can be buffered with a secondbuffer supply air 262B from the second circuit 276. In other words, thefirst circuit 274 can buffer a first portion(s) of the gas turbineengine 20, while the second circuit 276 can buffer a second, differentportion(s). Example components subject to low pressure requirementsinclude bearing compartments in low pressure regions of the gas turbineengine 20, such as front or rear bearing compartments. Examplecomponents subject to high pressure requirements include bearingcompartments in high pressure regions of the gas turbine engine 20, suchas mid-engine bearing compartments.

In this example, the first circuit 274 is similar to the buffer system60 of FIG. 3 and includes a first bleed air supply 264A, a second bleedair supply 266A and a valve 268A. The second circuit 276 includes afirst bleed air supply 264B, a second bleed air supply 266B, a valve268B and a conditioning device 280. In this non-limiting example, theconditioning device 280 cools the second buffer supply air 262B to anacceptable temperature for addressing higher pressure requirements. Theconditioning device could include an air-to-air heat exchanger, afuel-to-air heat exchanger, or any other suitable heater exchanger. Theconditioning device 280 could also be a device other than a heatexchanger.

The second bleed air supply 266A of the first circuit 274 can be commonto the first bleed air supply 264B of the second circuit 276. Thesesources can also be completely separate. In each of the first circuit274 and the second circuit 276, the second bleed air supplies 266A, 266Bare communicated as the buffer supply airs 262A, 262B for low powerconditions of the gas turbine engine 20 and the first bleed air supplies264A, 264B are communicated as the buffer supply airs 262A, 262B inresponse to high power conditions of the gas turbine engine 20. Examplelow power conditions include, but are not limited to, ground operation,ground idle and flight idle conditions. Example high power conditionsinclude, but are not limited to, takeoff, climb, and cruise conditions.It should be understood that other power conditions are alsocontemplated as within the scope of this disclosure.

In one exemplary embodiment, the valves 268A, 268B select andcommunicate the first bleed air supplies 264A, 264B (which arerelatively lower pressure bleed air supplies) as the buffer supply airs262A, 262B in response to identifying a high power condition of a gasturbine engine 20. The second bleed air supplies 266A, 266B (which arerelatively higher pressure bleed air supplies) are selected by thevalves 268A, 268B and communicated as the buffer supply airs 262A, 262Bin response to detecting a low power condition of the gas turbine engine20. Both the lower bleed air supplies and the higher bleed air suppliesare intended to maintain the same minimum pressure delta across thebearing compartment seals. Low power conditions require a higher stagepressurize source to contain the lubricant within the bearingcompartment, while high power conditions require a lower pressure stagesource. The buffer system 260 can use the lowest possible compressorstage to meet the pressure requirements in order to minimize supplytemperature and any performance impact to the gas turbine engine 20.

The buffer system 260 can also include a controller 270 in communicationwith the valves 268A, 268B for selectively switching between the firstbleed air supplies 264A, 264B and the second bleed air supplies 266A,266B. A single controller or multiple controllers could be utilized. Thecontroller 270 can also command operation of the conditioning device 280of the second circuit 276 for cooling the buffer supply air 262B.Alternatively, separate controllers can be used to control each of thefirst circuit 274, the second circuit 276 and the conditioning device280.

FIG. 6 illustrates another exemplary buffer system 360. Like the buffersystem 260, the example buffer system 360 is a two-circuit, multi-sourcebuffer system that includes at least a first circuit 374 and secondcircuit 376. Additional circuits could also be incorporated. Lowpressure requirements of the gas turbine engine 20 can be fed with afirst buffer supply air 362A from the first circuit 374, while highpressure requirements of the gas turbine engine 20 can be buffered witha second buffer supply air 362B from the second circuit 376. In otherwords, the first circuit 374 can buffer a first portion or portions ofthe gas turbine engine 20, while the second circuit 376 can buffer asecond, different portion or portions.

In the exemplary embodiment of FIG. 6, the first circuit 374 is similarto the buffer system 160. In this exemplary embodiment, the firstcircuit 374 includes an ejector 372 that can selectively mix a firstbleed air supply 364 having a first pressure with a second bleed airsupply 366 having a second, higher pressure to render a buffer supplyair 362A of an intermediate pressure. In one example, the ejector 372 isa variable area ejector that can be either actively or passivelycontrolled. An example variable area ejector is shown in FIG. 7, thefeatures of which are further discussed below. It should be understoodthat the ejector 372 could be used in the second circuit 376 or in bothcircuits 374, 376.

Augmentation of the first bleed air supply 364 prepares a buffer supplyair 362A at an adequate pressure and temperature to pressurize any lowpressure requirements of the gas turbine engine 20. The determination ofwhether or not to augment the first bleed air supply 364 with theejector 372 is based on a power condition of the gas turbine engine 20,or alternatively, is based on direct pressure measurement. Gas turbineengine power conditions can include low power conditions and high powerconditions. Example low power conditions include, but are not limitedto, ground operation, ground idle and flight idle conditions. Examplehigh power conditions include, but are not limited to, takeoff, climb,and cruise conditions. It should be understood that other powerconditions are also contemplated as within the scope of this disclosure.

In one example, the first bleed air supply 364 is augmented by theejector 372 in response to detecting a low power condition of the gasturbine engine 20 in order to communicate a buffer supply air 362Ahaving adequate pressurization. The amount of augmentation performed onthe first bleed air supply 364 can vary depending upon the type of powercondition that is detected and pressure requirements. For example, inone embodiment, the first bleed air supply 364 is not augmented by theejector 372 in response to detection of a high power condition of thegas turbine engine 20. In other words, the first bleed air supply 364can be communicated as the buffer supply air 362A without anyaugmentation in response to some power conditions.

The exemplary second circuit 376 of the buffer system 360 can include athird bleed air supply 365 (which may or may not be common to either thefirst bleed air supply 364 or the second bleed air supply 366 of thefirst circuit 374), a fourth bleed air supply 367 (which may or may notbe common to either of the first bleed air supply 364 or the secondbleed air supply 366 of the first circuit 374), a valve 368 and aconditioning device 380. The conditioning device 380 can cool the secondbuffer supply air 362B to an acceptable temperature for addressinghigher pressure requirements. The conditioning device 380 could includean air-to-air heat exchanger, a fuel-to-air heat exchanger, or any othersuitable heat exchanger, or an ejector.

In one example, the fourth bleed air supply 367 is communicated as thebuffer supply air 362B during low power conditions and the third bleedair supply 365 is communicated as the buffer supply air 362B during highpower conditions of the gas turbine engine 20.

The buffer system 360 can also include a controller 370 in communicationwith the ejector 372 and the valve 368 for selectively controlling thecommunication of the buffer supply airs 362A, 362B at an appropriatepressure and temperature. A single controller or multiple controllerscan be utilized. The controller 370 can also command operation of theconditioning device 380 of the second circuit 376 for cooling the buffersupply air 362B.

FIG. 7 illustrates the example ejector 372 of FIG. 6. It should beunderstood that the ejector 372 could be incorporated into one or moreof the buffer systems detailed above. For example, the ejector 372 couldbe used in place of the ejector 172 of FIG. 4. The ejector 372 can be avariable area ejector that is either passively or actively controlled.

The example ejector 372 includes a first inlet 101 for receiving thefirst bleed air supply 364 (of a relatively lower pressure), a secondinlet 103 for receiving the second bleed air supply 366 (of a relativelyhigher pressure), a mixing section 105, a diffuser section 107, and anozzle 109. The second bleed air supply 366 is communicated through thesecond inlet 103 and into the nozzle 109. The nozzle 109 reduces thepressure of the second bleed air supply 366 below the static pressure ofthe first bleed air supply 364 by forcing it through an orifice 113 thatcauses it to accelerate. The first bleed air supply 364 is drawn throughthe first inlet 101 by the static pressure differential between itselfand the accelerated second bleed air supply 366 and mixes with thesecond bleed air supply 366 in the mixing section 105 to render a buffersupply air 362A having an intermediate static pressure to the firstbleed air supply 364 and the second bleed air supply 366. The diffusersection 107 decelerates the buffer supply air 362A prior tocommunicating the buffer supply air 362A to a low pressure requirementof the gas turbine engine 20 so that the mixed flow static pressure isabove that of the first bleed air supply 364.

A plunger 111 can be movably positioned within the nozzle 109 to varythe orifice 113 of the nozzle 109. An actuator 115 is positioned to movethe plunger 111 in the direction D. Varying the position of the plunger111 within the nozzle 109 thereby controls the flow rate of the secondbleed air supply 366 into the mixing section 105. A controller, such asthe controller 370, can be programmed to selectively move the plunger111 to vary the exit area 113 of the nozzle 109.

Although the different examples have a specific component shown in theillustrations, embodiments of this disclosure are not limited to thoseparticular combinations. It is possible to use some of the components orfeatures from one of the examples in combination with features orcomponents from another one of the examples.

Furthermore, the foregoing description shall be interpreted asillustrative and not in any limiting sense. A worker of ordinary skillin the art would understand that certain modifications could come withinthe scope of this disclosure. For these reasons, the following claimsshould be studied to determine the true scope and content of thisdisclosure.

What is claimed is:
 1. A gas turbine engine, comprising: a buffer systemthat communicates buffer supply air to a portion of the gas turbineengine, wherein said buffer system includes: a first circuit thatselects between a first bleed air supply having a first pressure and asecond bleed air supply having a second pressure that is greater thansaid first pressure to render a first buffer supply air having anintermediate pressure compared to said first pressure and said secondpressure; and a second circuit that selects between a third bleed airsupply and a fourth bleed air supply to communicate a second buffersupply air.
 2. The gas turbine engine as recited in claim 1, whereineach of said first circuit and said second circuit include at least oneof an ejector and a valve.
 3. The gas turbine engine as recited in claim2, wherein said ejector is a variable area ejector that includes aplunger movably positioned within a nozzle to vary an orifice associatedwith said nozzle.
 4. The gas turbine engine as recited in claim 1,wherein one of said first circuit and said second circuit includes aconditioning device.
 5. The gas turbine engine as recited in claim 1,wherein said first buffer supply air is communicated to a componentsubject to a low pressure requirement of the gas turbine engine and saidsecond buffer supply air is communicated to a component subject to ahigh pressure requirement of the gas turbine engine.
 6. The gas turbineengine as recited in claim 1, wherein said portion includes at least abearing compartment of the gas turbine engine.
 7. The gas turbine engineas recited in claim 1, comprising at least one controller incommunication with each of said first circuit and said second circuit.8. The gas turbine engine as recited in claim 7, wherein said controllerselectively controls the communication of said first buffer supply airand said second buffer supply air.
 9. A gas turbine engine, comprising:a compressor section; a combustor in fluid communication with saidcompressor section; a turbine section in fluid communication with saidcombustor; a buffer system including a first circuit that supplies afirst buffer supply air and a second circuit that supplies a secondbuffer supply air; wherein said first circuit includes a first bleed airsupply, a second bleed air supply and at least one of an ejector and avalve; and wherein said second circuit includes a third bleed airsupply, a fourth bleed air supply and at least one of an ejector and avalve.
 10. The gas turbine engine as recited in claim 9, wherein one ofsaid first circuit and said second circuit includes a conditioningdevice.
 11. The gas turbine engine as recited in claim 9, wherein saidejector is a variable area ejector.
 12. The gas turbine engine asrecited in claim 9, wherein the gas turbine engine is a high bypassgeared aircraft engine having a bypass ratio of greater than about six(6).
 13. The gas turbine engine as recited in claim 9, wherein the gasturbine engine includes a low Fan Pressure Ratio of less than about1.45.
 14. The gas turbine engine as recited in claim 9, wherein saidfirst circuit includes one of said ejector and said valve and saidsecond circuit includes the other of said ejector and said valve.
 15. Amethod of cooling a portion of a gas turbine engine, comprising: mixinga first bleed air supply with a second bleed air supply to render afirst buffer supply air of an intermediate pressure to the first bleedair supply and the second bleed air supply; communicating the firstbuffer supply air to a component subject to a first pressure requirementof the gas turbine engine; and communicating a second buffer supply airto a component subject to a second pressure requirement of the gasturbine engine.
 16. The method as recited in claim 15, wherein thecomponent subject to the second pressure requirement is located separatefrom the component subject to the first pressure requirement.
 17. Themethod as recited in claim 15, comprising the step of: identifying apower condition of the gas turbine engine prior to the step of mixingand the steps of communicating.
 18. The method as recited in claim 17,comprising the step of: performing the step of mixing in response todetecting a low power condition during the step of identifying; orcommunicating the first bleed air supply as the first buffer supply airwithout performing the step of mixing in response to detecting a highpower condition during the step of identifying.
 19. The method asrecited in claim 18, comprising the step of: communicating a third bleedair supply as the second buffer supply air in response to a high powercondition of the gas turbine engine; or communicating a fourth bleed airsupply as the second buffer supply air in response to a low powercondition of the gas turbine engine.
 20. The method as recited in claim15, comprising the step of: cooling the second buffer supply air priorto the step of communicating the second buffer supply air to thecomponent subject to the second pressure requirement.